Airfoil structures in modern gas turbine engines, including those forming portions of rotating turbine blades and stationary nozzle vanes, are subjected to extremely high temperatures due to impingement thereon of hot combustion gas flow. In order to maintain acceptable mechanical properties in this harsh environment, metal blades and vanes are routinely cooled internally by air bled from a compressor portion of the engine. Since cooling air is not available to be mixed with fuel, ignited in the combustor and undergo work extraction in the primary gas flowpath of the turbine, cooling flow is treated as a parasitic loss in the engine operating cycle, it being desirable to keep such losses to a minimum.
Various schemes are employed to enhance cooling of an airfoil with a predetermined flow rate of cooling air so as to maintain an acceptable airfoil temperature profile. Such schemes include creating one or more flow passages within the airfoil to direct the cooling flow in an advantageous manner, for example by first directing the cooling flow to the hottest portion of the airfoil such as a leading edge. Additionally, internal sides of airfoil pressure and suction walls are often provided with obstructive surface features such as turbulator ribs, strips or pins which extend into the flow passage. By causing interruption in the thermal boundary layer proximate the walls, the cooling flow separates from and reattaches to the walls, increasing the convective heat transfer between the airfoil walls and the coolant flowing thereby, over that of a smooth wall condition. The size, quantity and orientation of turbulators on the pressure and suction walls with respect to the coolant flow are selected by those having skill in the art to tailor cooling within the constraints imposed by the geometry of the airfoil and the available coolant flow. Examples of turbulated passages in a turbine blade and a casting core for the manufacture thereof are disclosed in related U.S. Pat. Nos. 4,514,144 and 4,627,480 entitled "Angled Turbulence Promoter" granted to Lee on Apr. 30, 1985 and Dec. 9, 1986, respectively, and assigned to the same assignee as the present invention. While the introduction of turbulators generally increases convective cooling of the airfoil, cooling may suffer when blockage of the flow passages becomes excessive, for example due to the number and height of the turbulators.
When further heat transfer augmentation is required to provide acceptable mechanical properties in the airfoil after optimizing turbulator configuration on the airfoil walls, the volumetric flow rate of coolant may be increased and/or the source of the coolant may be changed to provide lower temperature air to the airfoil to increase the cooling thereof. In either case, such a change increases parasitic losses in the engine with a concomitant reduction in engine operating efficiency. In an existing engine design where airfoil cooling has been determined to be marginal or inadequate, the cost of modifying hardware to provide more or lower temperature cooling flow may be prohibitive. In this instance, blades and vanes could be replaced with components manufactured from a more suitable material, if available, or the existing hardware may be replaced more frequently than would otherwise be required if cooling were adequate.